Onboard supplemental power system at varying high altitudes

ABSTRACT

Systems and methods for supplementing a power system to achieve consistent operation at varying altitudes are disclosed herein. A hybrid power system comprising a single power source driving multiple generators may implement a power recovery turbine to drive a supercharger compressor, which may provide compressed air at increased altitudes. The supplemental power system disclosed herein provides necessary shaft horsepower at high altitudes to drive a generator and produce cooling air.

STATEMENT REGARDING FEDERALLY SPONSORED RESEARCH OR DEVELOPMENT

The U.S. Government has a paid-up license in this invention as providedfor by the terms of contract No. F33615-03-2367 awarded by USAF/AFMC.

BACKGROUND

1. Field

The present disclosures herein relate to aircraft engines, and morespecifically to systems and techniques for augmenting the powercapability of aircraft engines in high altitude environments.

2. Background

Ongoing development and growth in the area of onboard aircraftelectrical systems and electronic sub-systems has resulted in a desireto augment existing aircraft systems with supplemental electricalgenerating capability. However, standard aircraft production-designcharacteristics generally leave little, if any, room for significantelectrical or cooling air systems expansion or modification. Thus, it isdifficult for aircraft to accommodate post-production systems additions.The traditional approach of original equipment manufacturers withrespect to expanding on-board power and electrical generating capabilityusually leads to extensive and costly aircraft and/or enginemodifications.

Aircraft are often powered by gas turbine engines, which have a greatpower-to-weight ratio compared to internal combustion reciprocatingengines. Gas turbine engines are commonly considered to be“over-powered” at low altitudes, because of their high power-to-weightratio. However, at high altitude, when the air gets thinner,air-breathing internal combustion engines lose power. Even gas turbineengines can quickly become “under-powered” as an aircraft ascends.Unfortunately, power enhancement modifications to an aircraft engineoften require costly structural alterations to the airframe itself.Thus, in addition to the main engines, aircraft often utilize additionalsmall gas turbine engines that may be installed within the aircraft.These additional engines may generate electric power and providepressurized air for power requirements while the aircraft is on theground. Generally, these devices have their functions taken over inflight by the main engine. However, as electrical requirements forpassenger amenities and other electronic needs have increased, theseauxiliary power units have become correspondingly larger. In modernaircraft, auxiliary power units are often utilized in-flight. Althoughmany auxiliary engines are now overpowered at sea level, they generallyare only able to provide constant power up to altitudes of about 25,000ft. (“FL25”), and have diminishing power as the increases beyond that.Gas turbine engines cannot easily be made any larger, as the increase insize and weight would require significant structural modification to theairframe itself.

In short, modern aircraft including military aircraft, which have highrequirements for electrical power, suffer deficiencies when equippedwith gas turbine engines, because they lose power at high altitude butcannot compensate with increased size due to airframe structurallimitations. Thus, the in-flight power generating capability of aircraftis often significantly limited under prior art constructs. One result isthat there is not currently a gas turbine power system capable ofoperating at high altitude with the ability to maintain the increasingdemand for more horsepower to drive a generator and produce cooling airin sufficient quantity, without requiring significant modification toairframe structures.

SUMMARY

In one aspect of the present invention, a gas turbine power system foran aircraft includes a gas turbine engine having a sensor systemconfigured to measure the air mass flow through the engine and anexhaust nozzle having a variable opening responsive to the sensorsystem, a power recovery turbine coupled to the variable opening in thegas turbine engine, a first compressor driven by the power recoveryturbine and configured to deliver compressed air to the gas turbineengine, and a second compressor coupled to the gas turbine engine or thepower recovery turbine.

In another aspect of the present invention, a method of regulating thepower of a gas turbine power system installed on an aircraft includesmeasuring the air mass flow through a gas turbine engine having an airintake and an exhaust outlet, adjusting, as a function of the measuredair mass flow, a variable opening nozzle coupled to the exhaust outletof the gas turbine engine, directing exhaust from the gas turbine enginethrough the adjusted variable opening nozzle, driving a power recoveryturbine with the exhaust, driving a first compressor with the powerrecovery turbine and routing compressed air generated by the firstcompressor to the air inlet of the gas turbine engine, and driving asecond compressor with the gas turbine engine or the power recoveryturbine.

In another aspect of the present invention, a gas turbine power systemfor an aircraft includes means for measuring the air mass flow through agas turbine engine, means, responsive to the means for measuring, forvariably opening an exhaust nozzle coupled to the gas turbine engine,means, coupled to the exhaust nozzle, for driving a first compressor,means for delivering a first portion of compressed air from the firstcompressor to the gas turbine engine, and means, coupled to the gasturbine engine or the means for driving the first compressor, forfurther compressing a second portion of the compressed air and routingit to an air conditioning system.

It is understood that other embodiments of the specific teachings hereinwill become readily apparent to those skilled in the art from thefollowing detailed description, wherein it is shown and described onlyseveral embodiments of the teachings by way of illustration. As will berealized, the subject matter of the teachings herein is capable of otherand different embodiments and its several details are capable ofmodification in various other respects, all without departing from thespirit and scope of these teachings. Accordingly, the drawings anddetailed description are to be regarded as illustrative in nature andnot as restrictive.

BRIEF DESCRIPTION OF THE DRAWINGS

Aspects of the disclosures herein are illustrated by way of example, andnot by way of limitation, in the accompanying drawings, wherein:

FIG. 1 is a schematic illustrating an exemplary system layout design;

FIG. 2 is a schematic illustrating aspects of the first system layoutdesign illustrated in FIG. 1;

FIG. 3 is a schematic illustrating a first alternative system layoutdesign;

FIG. 4 is a schematic illustrating a second alternative system layoutdesign; and

FIG. 5 is a schematic illustrating a third alternative system layoutdesign.

DETAILED DESCRIPTION

The detailed description set forth below in connection with the appendeddrawings is intended as a description of various embodiments of theteachings herein and is not intended to represent the only embodimentsin which the teachings herein may be practiced. The term “exemplary”used throughout this disclosure means “serving as an example, instance,or illustration,” and should not necessarily be construed as preferredor advantageous over other embodiments. The detailed descriptionincludes specific details for the purpose of providing a thoroughunderstanding of the teachings. However, it will be apparent to thoseskilled in the art that the teachings herein may be practiced withoutthese specific details. In some instances, well-known structures anddevices are shown in block diagram form in order to avoid obscuring theconcepts of the teachings herein. Acronyms and other descriptiveterminology may be used merely for convenience and clarity and are notintended to limit the scope of the teachings herein. The term “coupled”is used throughout this disclosure to indicate structural or functionalcooperation between two components. In the case of structuralcooperation, the components may be connected directly to one another or,where appropriate in the context, connected indirectly to one another,e.g., through intervening or intermediary devices or other means. In thecase of functional cooperation, there may or may not be a physicalconnection between the two components.

In the following detailed description, various aspects of the teachingsherein will be described in the context of a gas turbine engine that maycomprise a commercially available off the shelf gas turbine engine.While these inventive aspects may be well suited for use with such anengine, those skilled in the art will readily appreciate that they arelikewise applicable for use in various other exhaust-producing aircraftengines. Accordingly, any reference to a gas turbine engine is intendedonly to illustrate various aspects of the disclosures herein, with theunderstanding that such aspects have a wide range of applications.

The teachings herein apply to aircraft gas turbine engine, to augmentits power capability so that it may produce sufficient power to run agenerator and cooling systems for power and cooling demands of theaircraft at high altitudes. In an exemplary embodiment, certainmodifications may be made to a gas turbine engine in order to increaseits power. For example, a supercharger may be utilized to boost thepressure of ambient air at high altitudes and deliver the pressurizedair to the gas turbine engine's air intake. Implementing thesupercharger would maintain atmospheric pressure at the air intake ofthe gas turbine engine, enabling the engine to provide sufficient powerto drive onboard electrical and cooling systems even as the aircraftflies at high altitudes.

Generally speaking, supercharging may be used to force compressed airinto a gas turbine engine to achieve improved engine performance andfuel efficiency. The supercharger may be driven by a power recoveryturbine, which in turn may be driven by exhaust gases from the gasturbine engine. The increase in air fed into the gas turbine engine bythe supercharger may increase combustion force and power. This increasemay compensate for thinner air at high altitudes, and prevent the enginefrom losing power as the aircraft climbs. The stabilized power producedby the gas turbine engine may be directed through an output shaft todrive rotors, compressors, ducted fans, or for any other intended usethe system designer may have for such power. In accordance with theteachings herein, one or more superchargers may be driven by the gasturbine engine exhaust, or coupled to the power output shaft of the gasturbine engine itself. These general concepts will be explained infurther detail below.

FIG. 1 is an schematic illustrating an exemplary supplemental powersystem installation that may be used, for example, on aircraft. Theexemplary supplemental power system may provide an aircraft withelectrical power and compressed air for continuous ground and airborneoperations. For example, it may provide for conditioning of the aircraftcockpit and passenger/cargo cabin areas, main engine starting and otherelectrical power requirements, at sea level and at high altitudes up toand beyond 40,000 feet (“FL40”). The exemplary supplemental power systemmay provide constant mass flow, variable inlet volumetric flow, and avariable compression ratio. That is, as the altitude of an aircraftincreases, the exemplary supplemental power system may increasinglycompress additional amounts of air, and feed this additional volume ofcompressed air to the air intake of the aircraft's gas turbine engine.As the aircraft increases altitude and encounters thinner air, thevariably increased volumetric flow of air through the gas turbine enginewill enable the gas turbine engine to experience a constant flow ofpounds of air (“air mass flow”). In other words, larger amounts ofcompressed air at higher altitudes, where the air is thinner (i.e. hassmaller mass), will approximate the air mass flow of smaller amounts ofuncompressed (ambient pressure) air at lower altitudes, where the air isthicker (i.e. has larger mass). Thus, a gas turbine engine processing agreater volumetric flow of air at high altitudes may generate the samelevel of power as if it were processing a smaller volumetric flow of airat seal level, because it would actually be processing the same “airmass flow” in each case.

In an exemplary embodiment, a gas turbine engine 100 may include acompressor, a combustion area and a turbine. The compressor (not shown)may be located near the air intake of the gas turbine engine 100 andraise the pressure of incoming air to produce pressurized air. The highpressure air may then enter the combustion area (not shown), where fuelinjectors may inject a stream of fuel, such as jet fuel in the case ofan aircraft. In the combustion area, the air and fuel are mixed. Thecombustion area burns the fuel and produces exhaust, which may provideboth power for use on board the aircraft, and thrust to cause theaircraft to move. The turbine (not shown) may be used to transmit powerto other systems on the aircraft, by driving an output shaft 101. Theturbine may include a set of vanes placed in the exhaust stream, that“catch” the exhaust and cause the vanes to spin, like a windmill. Thevanes may be attached to the output shaft, which will spin as the vanesin the exhaust stream spin. The output shaft 101 may then be used todrive other systems on the aircraft. In this manner, the turbineextracts energy from the high pressure, high velocity exhaust that flowsfrom the combustion chamber, and transmits the extracted energy throughthe output shaft 101. This energy may be used to provide power forelectrical, cooling, and other systems on board the aircraft, that maybe coupled directly or indirectly to the output shaft 101.

As mentioned above, in addition to producing power the gas turbineengine 100 may provide thrust for causing the aircraft to move forward.A nozzle may be formed at the exhaust end of the gas turbine engine, togenerate a high speed jet of exhaust gas. This high speed exhaust jetmay provide thrust that causes the aircraft to move forward. Therefore,the gas turbine engine 100 may provide both thrust to move the aircraftforward, and additional power for driving various electrical and coolingsystems on board the aircraft. The gas turbine engine 100 may comprise acommercial off the shelf (“COTS”) engine, as selected by a systemdesigner. For example, the gas turbine engine 100 may comprise a Pratt &Whitney PW127G turbine engine, capable of producing approximately2600–2800 shaft horsepower (“shp”) at sea level static, standard dayconditions. This is only one example of a gas turbine engine that may beutilized in the exemplary embodiment, and should not be read to limitthe teachings herein. Those skilled in the art will recognize that anyof a number of different engines may be used in conjunction with theteachings herein.

In accordance with the general principles disclosed herein, the gasturbine engine 100 may receive ambient air 102 at its air intake 128.The ambient air may be compressed, mixed with fuel, and combusted in thegas turbine engine 100, as explained above. The exhaust gases producedby the combusted fuel-air mixture may be used to rotate the output shaft101. The output shaft 101 of the gas turbine engine 100 may drive apower generator 106 that produces power as indicated at arrow 107.Alternatively, the power generator 106 may be driven by a separate,power recovery turbine, as will be explained in further detail below.The power generator 106 may comprise a 1000 kilivolt ampere (“KVA”)generator, or any other generator that may be required to supportvarious electrical, cooling, and other systems on board an aircraft tobe fitted with the exemplary supplemental power system.

To enable it to provide sufficient power to drive generator 106 at highaltitude, the gas turbine engine 100 may be supercharged. A powerrecovery turbine 112 coupled to the output of the gas turbine engine100, either directly or indirectly, may be used to drive a superchargercompressor 108. The power recovery turbine 112 may be coupled to avariable area nozzle (“VAN”) 114 at the exhaust outlet of the gasturbine engine 100. The VAN 114 may be opened or closed to change thesize of the nozzle through which exhaust from the gas turbine engine 100may be discharged. As the VAN 114 is opened, the nozzle area becomeslarger, allowing exhaust to escape more easily. As the VAN 114 isclosed, the nozzle area becomes smaller, partially blocking the path ofexhaust from the gas turbine engine 100. Thus, closing or partiallyclosing the VAN 114 will increase backpressure at the exhaust outlet ofthe gas turbine engine 100.

An exhaust duct 116, coupled to the exhaust outlet of the gas turbineengine 100, may include a bypass duct 118 for pressure relief. When thegas turbine engine 100 does not need to be supercharged in order tosupport the aircraft's on board systems, such as at lower altitudes,exhaust from the gas turbine engine 100 may pass through bypass duct118. However, when additional power is required, such as at higheraltitudes, exhaust may be directed past the bypass duct 118 and throughthe power recovery turbine 112. This re-direction of exhaust may beaccomplished, for example, by closing or partially closing the VAN 114and increasing the backpressure at the exhaust outlet of the gas turbineengine 100. The details regarding how this may be accomplished will beexplained in further detail below. In any case, as exhaust flows throughthe power recovery turbine 112, it may rotate the power recovery turbine112 having an output shaft 110 attached thereto. The output shaft 110may be used to drive the supercharger compressor 108, which maysupercharge the gas turbine engine 100. Supercharging the gas turbineengine 100 involves feeding additional volumes of air to the air intake128 of the gas turbine engine 100. This may sustain the level of airmass flow through the gas turbine engine and thus the power produced bythe gas turbine engine 100, even at the aircraft climbs to higheraltitudes where the ambient air is thinner.

The supercharger compressor 108 may be either axial flow or radial, andmay be a low-pressure compressor (LPC) which can be sized to produceadditional volumes of air to supply air flowing either directly to airconditioning equipment or to another compressor for other coolingsystems on board the aircraft. The supercharger compressor 108 mayinclude inlet guide vanes (“IGV”) 109 to regulate the amount of ambientair 120 air that enters. The supercharger compressor 108 may compressthe ambient air 120, which results in compressed air 122. The compressedair 122 may pass through an intercooler 124, which may cool thecompressed air 122 as well as ambient air 126. This cooled air may beused to supplement the air that is received by the gas turbine engine100, as well as to support air conditioning and other cooling systemsthat may be on board the aircraft. Both applications will be explainedin further detail below.

If the intercooler 124 generates cooled, compressed air in a quantitythat exceeds system requirements imposed by the generator 106 and theair conditioning system, the excess air may be released at a bypass duct132. However, the majority of cooled, compressed air produced by theintercooler 124 may be used for both the gas turbine engine 100 and theair conditioning or other cooling systems. Cooling the compressed air122 before it reaches the air intake 128 of the gas turbine engine 100may preserve or increase the air mass flow of this compressed air 122.Generally speaking, when air is compressed, such as by the superchargercompressor 108, its temperature rises. If some or all of the air takeninto a gas turbine engine has been supercharged, the benefit ofsupercharging (i.e. greater mass flow) may be reduced by thistemperature rise. Thus, a cooling device such as the intercooler 124 maybe employed to reduce the temperature of the supercharged air andpreserve the greater air mass flow. Cooled air from the intercooler 124may then be directed to the air intake 128 of the gas turbine engine100. The gas turbine engine 100 may thus receive both the ambient air102 and the additional cooled air from the intercooler 124. Thisadditional amount of air from the intercooler 124 may increase pressureat the air intake 128, until it approximates the backpressure at theexhaust outlet of the gas turbine engine 100. Thus, if the backpressurehad previously been increased, such as by closing or partially closingthe VAN 114, the additional air supplied to the gas turbine engine 100via supercharging may cause an equivalent increase in pressure at theair intake 128.

In accordance with the teachings above, if the backpressure is adjusted,such as to approximate the typical backpressure that would be present atsea level conditions, the exemplary supplemental power system may beengaged as described above to increase the pressure at the air intake128 so that it also approximates the typical air intake pressure thatwould be present at sea level conditions. By repeating this process ofcontrolling the VAN 114 to increase backpressure at the exhaust 116,then driving the supercharger compressor to feed more air into the airintake 128 and raise the pressure at the air intake 128, the gas turbineengine 100 will be able to process a consistent air mass flow andproduce a consistent level of power, even as altitude increases and theambient air 102 becomes thinner. Further, this consistent air mass flowmay be controlled such that it is approximately what the gas turbineengine 100 would experience at sea level, causing the gas turbine engineto operate at full-powered sea level conditions, even at high altitudes.

In addition to supplementing the air intake of the gas turbine engine100, cooled air from the intercooler 118 may be directed to a loadcompressor 104 that supports air conditioning and other cooling systemson board the aircraft. The load compressor 104 may produce compressed,cooled air 130 that may be used, for example, in on-board airconditioning, component cooling, or other types of conditioning systemsthat may be on board the aircraft. In an exemplary embodiment, thecompressed, cooled air 130 may be approximately 800 pounds per minute,at 50 pounds per square inch absolute (“psia”). However, it will berecognized by those skilled in the art that exact specifications may bealtered in accordance with the present teachings and tailored to fit therequirements of cooling systems on board various aircraft, as necessary.

The effect of supplying the gas turbine engine 100 with additional airfrom supercharger compressor 108 as described above, as well asmechanisms for controlling the amount of additional air, will now beexplained. As illustrated in FIG. 1 and described in reference thereto,the exemplary supplemental power system may utilize the residual powerobtained from captured gas turbine engine exhaust to compress ambientair. The compressed air results in an increased volume of air, which maybe used to preserve the air mass flow processed by the gas turbineengine 100 at increased altitudes, for example, where ambient air isthinner. This additional air may compensate for the lower air-weightvolume ratios that occur at high altitude, which would otherwise reducethe power capabilities of gas turbine engine 100 when an aircraft fliesat higher altitudes. Preserving the air mass flow as altitude increasesmay enable the gas turbine engine 100 to generate a consistent level ofpower, even as the aircraft climbs. The air mass flow may be preservedby controlling the compressed air that is fed to the gas turbine engine100, such that the gas turbine engine 100 receives air at a weight equalto that which it would receive at sea level.

A control system 134 may regulate the amount of compressed air that isfed to the air intake 128 of the gas turbine engine 100, such that itreceives an approximately consistent air mass flow between sea level andhigher altitudes. The control system 134 may rely on a buildup ofbackpressure in the gas turbine engine 100 in order to produce sea leveloutput even at high altitudes up to at least 40,000 feet (“FL40”).Pressure may be monitored by one or more sensors, such as an intakesensor 138 located at the air intake 128 of the gas turbine engine 100,and an exhaust sensor 140 located in the exhaust duct 116 adjacent tothe variable area nozzle 114. The intake sensor 138 may include aplurality of sensors, and the exhaust sensor 140 may also include aplurality of sensors. The sensors may be pressure sensors or otherappropriate sensors for measuring or determining the pressure at variousareas within and around the gas turbine engine 100 in order to measurethe mass air flow through the gas turbine engine 100. The control system134 may receive input from the pressure sensors 138 and 140, as well asother input signals that will be described below, and correspondinglycontrol the VAN 114 to regulate backpressure at the exhaust outlet ofthe gas turbine engine 100.

The control system 134 may operate in conjunction with an aircraft'sFull Authority Digital Engine Control (“FADEC”) 136 and a cooling airdemand regulating system 138. The FADEC 136 may control the outputs ofthe gas turbine engine 100 and power recovery turbine 112 as theaircraft is climbing through to FL40. The FADEC may, after receivingreadings from an altimeter or other altitude sensing device, providealtitude information to the control system 134. The FADEC may alsoprovide information to the control system 134 regarding the powerrequirements of various systems on board the aircraft any a given pointin time. By knowing the altitude or the power requirements, controlsystem 134 may determine the power output requirement of generator 106,and in turn determine the amount of power that must be generated by thegas turbine engine 100. This information may be used by the controlsystem 134 to control the supercharger 108 such that the gas turbineengine 100 is sufficiently supercharged to generate the necessary amountof power. The cooling air demand regulating system 139 may provideinformation to the control system 134 regarding the aircraft's demandfor cooling air at a given time. This information may also be used bythe control system 134 to control the supercharger 108 such that theload compressor 104 receives a sufficient amount of compressed air tosupply the air conditioning and other cooling systems that may be onboard the aircraft.

The procedure by which the control system 134 may control thesupercharger 108 will now be explained in further detail. The controlsystem 134 may be designed to advantageously use the “balanced” designof a gas turbine engine such as gas turbine engine 100. Gas turbineengines are typically balanced by design, such that they operate withthe same atmospheric pressure at the engine exhaust as at the airintake. Thus, as altitude increases and the ambient air pressuredecreases, the control system 134 may adjust the backpressure of gasturbine engine 100 to compensate for the pressure decrease.Specifically, the control system 134 may cause the VAN 114 to close orpartially close, which will reduce the cross section of the exhaustoutlet of the power recovery turbine 100 and cause an increase inbackpressure, accordingly. The backpressure may be increased to thelevel of backpressure that is normally experienced at sea level, byclosing the VAN 114 an appropriate amount. This amount may bepre-determined based on altitude and pressure values, and programmedinto the control system 134. The control system 134 may be designed toproduce the increase in backpressure at a threshold altitude or at aseries of threshold altitudes, which may be determined by the systemdesigner. The increased backpressure may also be initiated based onreadings received from pressure sensors 138 and 140. The control systemmay utilize algorithms to determine the ratio of altitude to pressure,and activate the VAN 114 at certain threshold ratios. Appropriatetiming, based on altitude and pressure ratios, may be determined by asystem designer who is skilled in the art, and programmed in the controlsystem 134.

Once the VAN 114 is activated, the increased backpressure at the exhaustof the gas turbine engine 100 will result in a pressure drop between theexhaust outlet of the gas turbine engine 100 and the power recoveryturbine 112. This pressure differential will create a vacuum, causingexhaust from the gas turbine engine 100 to flow through and activate thepower recovery turbine 112. Driven by the power recovery turbine 112,the supercharger compressor 108 may augment air pressure at the airintake 128 of the gas turbine engine 100, such that it becomesequivalent to the increased backpressure at the exhaust outlet of thegas turbine engine 100. This procedure of increasing backpressure to sealevel conditions, which instigates an increase in pressure at the airintake 128 such that the air intake 128 is also at sea level conditions,may enable the gas turbine engine 100 to receive the same constant flowof air mass flow per minute that it was designed to receive at sealevel. It is to be understood that the gas turbine engine 100 may be anysuitable gas turbine engine, and that the control system 134 may beprogrammed to achieve air mass flow conditions appropriate for whatevergas turbine engine is selected for a particular application.

Under control of the control system 134, the backpressure of the turbineengine 100 may be incrementally increased to approximate typical sealevel backpressure conditions as the aircraft gains altitude.Controlling the VAN 114 to create sea level backpressure will in turncontrol the amount of exhaust that is directed to the power recoveryturbine. This may be accomplished under control of the control system134, described above. The increase in backpressure at each increment mayinstigate or augment compressor and turbine stages of superchargercompressor 108, to supercharge the gas turbine engine 100. In thismanner, i.e. by increasing the amount of exhaust the power recoveryturbine 112 it receives from gas turbine engine 100, the power recoveryturbine 112 may be controlled. The control system 134 may variably applythe incremental compressor and turbine stages of the superchargercompressor 108 to the gas turbine engine 100, adding more and more airvolumetrically while producing additional power by use of the powerrecovery turbine 112. When power recovery turbine 112 is not in use,such as at sea level, it can be removed from operation of the overallsystem such as by the bypass duct 118.

Actuating the VAN 114 from full open, to closure in stages, may be theprincipal means by which power for power recovery turbine 112 iscontrolled. Closing the VAN 114 produces greater backpressure and thusmore power; opening the VAN 114 reduces backpressure and the resultantpower. As an aircraft climbs and altitude increases, the VAN 114 mayprogressively close, which would impose increased backpressure orresistance to the exhaust gas flow from the engine. This in turn wouldcause a pressure drop from the engine exhaust outlet to the powerrecovery turbine exit. As explained above, the pressure drop wouldcreate a vacuum effect, which would draw exhaust through the powerrecovery turbine 112. When the IGV 109 are open or partially open, powerrecovery turbine 112 may drive supercharger compressor 108 to generateadditional air for the gas turbine engine's air intake 128. Byappropriate scheduling according to altitude and air mass flowconditions, algorithms for which may be programmed into control system134, the pressure at the engine exit can be maintained at sea levelconditions or incrementally adjusted to sea level conditions, providingincremental pressure drops for the power recovery turbine 112 to produceincrementally larger amounts of power from the remainder of the energyavailable at a particular altitude.

In another exemplary embodiment, the power recovery turbine 112 maycomprise a two-stage design. For a two-stage power recovery turbine,both stage nozzles may be VANs for maintaining more favorable pressuredrops through the stages and, thus, better overall turbine efficiency.The two VANs may be coupled or individually controlled. Input pressureat high altitude may be sensed by a transducer and sent to the controlsystem 134. The control system 134 may then open or close the variablearea nozzles by pre-programmed amounts, selected as a function ofaltitude. Because the goal is to keep the engine inlet conditionssimilar or equivalent to sea level conditions, control logic of thecontrol system 134 may include the coupling effect of the power recoveryturbine 112 and its influence on the gas turbine engine's backpressureand power.

In an exemplary embodiment of the control system 134, control algorithmsbased upon pre-determined altitude conditions may be programmed in thecontrol logic to adjust VAN 114 (or a combination of two VANs) inrelation to altitude. The control logic may be implemented according tothe algorithms in order to engage supercharger compressor 108sufficiently to maintain sea level inlet conditions of gas turbineengine 100. An example of such conditions for the control logic isprovided below. However, it is to be understood that these conditionsand instructions may be modified as appropriate to ensure satisfactoryoperations of the system.

In the exemplary scenario, at sea level altitude the IGV 109 may befully closed, and first and second variable area nozzles may be fullyopened. The gas turbine engine 100 and the load compressor 104 may befed through the bypass duct 118 as the gas turbine engine 100 operatesat sea level conditions. At 10,000 feet, the IGV 109 may remain fullyclosed. The first VAN may be partially closed, keeping sea levelconditions at the gas turbine engine exit. The power recovery turbine112 produces partial power, and the gas turbine engine 100 operates atnearly full load. At approximately 15,000 to 20,000 feet, the IGV 109may open partially. The first VAN may close further, and the second VANmay also partially close, creating sea level conditions at the engineexit. The supercharger compressor 108 may begin to operate, circulatinggas through the intercooler 142 to the gas turbine engine 100 and theload compressor 104. The power recovery turbine 112 produces partialpower. At 20,000 feet, the IGV 109 may fully open. The first VAN may besufficiently closed to produce sea level conditions at the gas turbineengine exit, and the system may be fully operational at this point.Between 20,000 and 40,000 feet, the system may be fully operational.During this altitude range, the second VAN may progressively close tomaintain the optimum pressure drop through the power recovery turbine112. Again, it is to be understood that the conditions and instructionsdescribed above may be modified as appropriate to ensure satisfactoryoperations of the system, and that the teachings of control system 134herein are not to be limited to the exemplary control logic instructionsprovided.

The exemplary control system embodiment described above may beimplemented in a variety of system layout designs. As will be recognizedby those skilled in the art, changes may be made to the particularexamples provided in the foregoing descriptions when constructing asystem according to the teachings herein. For example, the superchargercompressor 108 may be downsized to need, producing only sufficient airto supercharge the engine itself. Alternatively, it may be upsized whereunabsorbed power produced by an added power recovery turbine 112 isre-directed to the gas turbine engine's primary power output shaft bycoupling the power recovery turbine 112 to the gas turbine engine 100itself. The power recovery turbine 112 and the supercharger compressor108 may be allowed to “wind mill” with only a minor power penalty.Alternatively, the power recovery turbine 112 may be uncoupled by meansof a clutch 117 enabling it to uncouple from the gas turbine engine 100at low altitudes. When the power recovery turbine 112 is not in use, itcan also be bypassed by ducting such as exhaust duct 118. Various systemlayout designs may utilize a gas turbine engine in a turbo-shaftconfiguration to provide power required for the generator 106 and/orload compressor 104. The gas turbine engine 100, a turbine engine, maybe either specifically designed for a particular application, or maycomprise a COTS turbine engine. The power recovery turbine 112 may be alow pressure turbine, the supercharger compressor 108 may be a lowpressure compressor, and the load compressor 104 may be a high pressurecompressor. These turbines and compressors may be arranged in variousconfigurations with an intercooler and electrical generator to meetspecific systems application requirements at varying high altitudes.

FIG. 2 is a schematic illustrating aspects of the system layout designillustrated in FIG. 1. A generator 200 and a high pressure compressor(“HPC”) 202 may be driven by a turbo shaft gas turbine engine 204. Agearbox 206, included with the gas turbine engine 204, may be used toprovide optimum revolutions per minute (“rpm”) for either the generator200 or the HPC 202. A supercharger low pressure compressor (“LPC”) 208may be driven on a separate shaft by a power recovery low pressureturbine (“LPT”) 210. The LPT 210 may receive full exhaust from the gasturbine engine 204. An intercooler 212 may provide cooler inlettemperature air for the gas turbine engine 204, thereby increasing itspower capability and efficiency. Air may enter the LPC 208, pass throughthe intercooler 212 and then split to supply both the gas turbine engine204 and the HPC 202. Compressed air from the HPC 202 may be dischargedto air conditioning equipment 214. Air leaving the gas turbine engine204 may pass through the LPT 210, which in turn may drive thesupercharging LPC 208. The LPT exhaust gas is then discharged toambient, as illustrated at 216.

FIG. 3 is a schematic illustrating a first alternative system layoutdesign. In this configuration, a generator 300, LPC 302, LPT 304 and HPC306 may all be on a low-pressure spool and driven by the turbine engineexhaust gas from a gas turbine engine 308. The gas turbine engine 308may supply the LPT 304 with high pressure and high temperature exhaustgas. Thus, the LPT may provide a moderate to high load driving thegenerator 300 and the two compressors 302 and 306. LPT exhaust gas maybe discharged to ambient, as illustrated at 310. An intercooler 312 maycool air produced by the LPC 302. Cooled air from the intercooler 310may then split to produce air intake for the gas turbine engine 308, andcompressed air, from the HPC 306, for air conditioning equipment 314.

FIG. 4 is a schematic illustrating a second alternative system layoutdesign. In this configuration, a generator 400, LPC 402 and LPT 404 mayagain be on a low-pressure spool and driven by the turbine engineexhaust gas from gas turbine engine 406. However, a HPC 408 may be onthe gas turbine engine shaft. Gas from a gas turbine engine 406 maydrive the generator 400 and the LPC 402 on the low pressure spool. Airfrom the LPC 402 may pass through an intercooler 410 and then split tosupply both the gas turbine engine 406 and the HPC 408. Compressed airfrom the HPC 408 may be discharged to air conditioning equipment 412.Air leaving the gas turbine engine 406 may pass through the LPT 404,which in turn may drive the supercharging LPC 402. LPT exhaust gas isthen discharged to ambient, as illustrated at 414.

FIG. 5 is a schematic illustrating a third alternative system layoutdesign. In this configuration, a generator 500 may be placed on theshaft of a gas turbine engine 502. LPT 504 may drive HPC 506 and LPC 508on the low pressure spool. Air from the LPC 508 may pass through anintercooler 510 and then split to supply both the gas turbine engine 502and the HPC 506. Compressed air from the HPC 506 may be discharged toair conditioning equipment 512. Air leaving the gas turbine engine 502may pass through the LPT 504, which in turn may drive the superchargingLPC 508. LPT exhaust gas is then discharged to ambient, as illustratedat 514.

The previous description of the disclosed embodiments is provided toenable any person skilled in the art to make or use the teachingsherein. Various modifications to these embodiments will be readilyapparent to those skilled in the art, and the generic principles definedherein may be applied to other embodiments without departing from thespirit or scope of the teachings disclosed herein. Thus, the scope ofthe disclosures herein is not intended to be limited to the embodimentsshown and described, but is to be accorded the widest scope consistentwith the general principles and novel features disclosed herein.

1. A gas turbine power system for an aircraft, comprising: a gas turbineengine having a sensor system configured to measure the air mass flowthrough the engine, and an exhaust nozzle having a variable openingresponsive to the sensor system; a power recovery turbine coupled to thevariable opening in the gas turbine engine; a first compressor driven bythe power recovery turbine, and configured to deliver compressed air tothe gas turbine engine; a second compressor coupled to the gas turbineengine or the power recovery turbine; and an exhaust duct coupledbetween the variable opening and the power recovery turbine, wherein thesensor system comprises a first pressure sensor located adjacent to anair intake of the gas turbine engine, and a second pressure sensorlocated in the exhaust duct adjacent to the variable opening.
 2. The gasturbine power system of claim 1, further comprising a control systemconfigured to: receive first and second input signals from the first andsecond pressure sensors; calculate the air mass flow through the gasturbine engine as a function of the first and second input signals;receive a third input signal indicative of the altitude of the aircraft;and adjust, as a function of the calculated air mass flow and thealtitude, the opening of the exhaust nozzle.
 3. The gas turbine powersystem of claim 2 further comprising inlet guide vanes coupled to thefirst compressor and configured to control the amount of air that entersthe first compressor, wherein the control system is further configuredto adjust, as a function of the calculated air mass flow and thealtitude, the position of the inlet guide vanes.
 4. The gas turbinepower system of claim 1 wherein the exhaust duct comprises a bypass ductconfigured to release exhaust from the gas turbine engine that is notdirected into the power recovery turbine.
 5. A gas turbine power systemfor an aircraft, comprising: a gas turbine engine having a sensor systemconfigured to measure the air mass flow through the engine, and anexhaust nozzle having a variable opening responsive to the sensorsystem; a power recovery turbine coupled to the variable opening in thegas turbine engine; a first compressor driven by the power recoveryturbine, and configured to deliver compressed air to the gas turbineengine; a second compressor coupled to the gas turbine engine or thepower recovery turbine; and an air intercooler coupled between the firstcompressor and an air intake of the gas turbine engine.
 6. The gasturbine power system of claim 5 further comprising an air conditioningsystem coupled to the second compressor.
 7. A gas turbine power systemfor an aircraft, comprising: a gas turbine engine having a sensor systemconfigured to measure the air mass flow through the engine, and anexhaust nozzle having a variable opening responsive to the sensorsystem; a power recovery turbine coupled to the variable opening in thegas turbine engine; a first compressor driven by the power recoveryturbine, and configured to deliver compressed air to the gas turbineengine; a second compressor coupled to the gas turbine engine or thepower recovery turbine; and a clutch coupled between, and configured toengage and disengage, the gas turbine engine and the power recoveryturbine.
 8. A method of regulating the power of a gas turbine powersystem installed on an aircraft, the method comprising: measuring theair mass flow through a gas turbine engine having an air intake and anexhaust outlet; adjusting, as a function of the measured air mass flow,a variable opening nozzle coupled to the exhaust outlet of the gasturbine engine to approximate sea level back pressure; directing exhaustfrom the gas turbine engine through the adjusted variable openingnozzle; driving a power recovery turbine with the exhaust; driving afirst compressor with the power recovery turbine and routing compressedair generated by the first compressor to the air inlet of the gasturbine engine; and driving a second compressor with the gas turbineengine or the power recovery turbine.
 9. A method of regulating thepower of a gas turbine power system installed on an aircraft, the methodcomprising: measuring the air mass flow through a gas turbine enginehaving an air intake and an exhaust outlet; adjusting, as a function ofthe measured air mass flow, a variable opening nozzle coupled to theexhaust outlet of the gas turbine engine; directing exhaust from the gasturbine engine through the adjusted variable opening nozzle; driving apower recovery turbine with the exhaust; driving a first compressor withthe power recovery turbine and routing compressed air generated by thefirst compressor to the air inlet of the gas turbine engine; and drivinga second compressor with the gas turbine engine or the power recoveryturbine; and wherein the adjustment of the variable opening nozzle isalso a function of the altitude of the aircraft.
 10. The method of claim9 wherein the adjustment of the variable opening nozzle comprises atleast partially closing the nozzle as the altitude increases.
 11. Themethod of claim 8 further comprising adjusting, as a function of themeasured air mass flow, inlet guide vanes coupled to the firstcompressor.
 12. The method of claim 11 wherein the adjustment of theinlet guide vanes is also a function of the altitude of the aircraft.13. The method of claim 12 wherein The adjustment of the inlet guidevanes comprises at least partially opening the inlet guide vanes as thealtitude increases.
 14. The method of claim 8 wherein the measuring ofair mass flow through the gas turbine engine comprises measuring apressure differential between the air intake and the exhaust outlet ofthe gas turbine engine.
 15. The method of claim 14 wherein the measuringof the pressure differential comprises: obtaining a first pressuremeasurement from a first pressure sensor located adjacent to the airintake of the gas turbine engine; obtaining a second pressuremeasurement from a second pressure sensor located adjacent to thevariable opening nozzle; and calculating the diffidence between thefirst and second pressure measurements.
 16. The method of claim 8further comprising directing exhaust from the gas turbine engine that isnot passed through the variable opening nozzle through a bypass outletand away from the power recovery turbine.
 17. The method of claim 8further comprising cooling air from the first compressor before it isrouted to the gas turbine engine.
 18. The method of claim 17 furthercomprising directing a portion of the compressed, cooled air to thesecond compressor.
 19. The method of claim 18 further comprisingdirecting air from the second compressor to an air conditioning system.20. A gas turbine power system for an aircraft, comprising: means formeasuring the air mass flow through a gas turbine engine; means,responsive to the means for measuring, for variably opening an exhaustnozzle coupled to the gas turbine engine; means, coupled to the exhaustnozzle, for driving a first compressor; means for delivering a firstportion of compressed air from the first compressor to the gas turbineengine; and means, coupled to the gas turbine engine or the means fordriving the first compressor, for further compressing a second portionof the compressed air and routing it to an air conditioning system.